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Shock-Induced Separation of a Laminar Boundary Layer in Supersonic Flow Past a Convex Corner

Published online by Cambridge University Press:  07 June 2016

N. Curle
N. Curle*
Affiliation:
The University, Southampton
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Summary

This paper presents a theoretical investigation of the interactions which occur when a laminar boundary layer on a flat surface encounters an expansive corner followed by a shock or other compressive agency. The method consists basically of two equations; the first relates deflections in the external streamlines to pressure changes and the second relates pressure changes to thickening or thinning of the boundary layer. For simplicity it is assumed that the pressure changes and streamline deflections are small.

A comparison is made with the only available experimental results, under conditions which are far outside the range of applicability of the theory. In these circumstances the agreement between the predicted and experimental pressure distributions is reasonable.

Type
Research Article
Information
Aeronautical Quarterly , Volume 16 , Issue 1 , February 1965 , pp. 33 - 41
Copyright
Copyright © Royal Aeronautical Society. 1965

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References

1. Robinson, M. L. Unpublished experimental results.Google Scholar
2. Curle, N. The Effects of Heat Transfer on Laminar Boundary-Layer Separation in Supersonic Flow. Aeronautical Quarterly, Vol. XII p. 309, November 1961.CrossRefGoogle Scholar
3. Thwaites, B. Approximate Calculation of the Laminar Boundary Layer. Aeronautical Quarterly, Vol. I p. 245, November 1949.CrossRefGoogle Scholar